Unitized hybrid rocket system

ABSTRACT

The hybrid rocket system of this invention is characterized by use of an oxidizer tank having a cylindrical midsection surrounded by a skirt and bonded thereto by a layer of elastomeric adhesive. The skirt outer surface is in turn adhesively secured to a spacecraft inner surface. An elongated solid-fuel motor case is mechanically rigidly secured to a central rear surface of the tank, and the case terminates in a throat and nozzle. The elastomeric-adhesive bonding of tank to skirt, and rigid adhesion of skirt to spacecraft forms the sole support for the rocket system, and separate support for the motor case is not required.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a National Stage application of International PCTapplication No. PCT/US2004/009694, filed on Mar. 29, 2004, which claimsthe priority to U.S. Provisional application No. 60/458,296, filed onMar. 28, 2003.

BACKGROUND OF THE INVENTION

This invention relates to a hybrid rocket system useful, for example, topropel a winged suborbital spacecraft, but not limited to thatapplication. In contrast to solid-fuel and liquid-fuel rockets, a hybridrocket motor uses both types of fuel. That is, a fluid oxidizer such asnitrous oxide (N₂O) is used to burn a solid fuel. The oxidizer is heldin a pressurized tank, and the solid fuel (such as HTPB, orhydroxyl-terminated polybutadiene) is cast on the inner walls of ahollow, and typically tubular combustion-chamber housing or motorextending rearwardly from the oxidizer tank, and terminating in a throatand nozzle. A pilot-controlled valve admits oxidizer to the housing, andan igniter (such as a spark or flame type) initiates combustion.

This invention is directed to two significant improvements. First, to asimplified method of securing and mounting the rocket to a spacecraftfuselage or associated structure; and second, to an integrated motorconstruction which sharply limits possible leakage paths for improvedreliability and safety.

SUMMARY OF THE INVENTION

The hybrid rocket system of this invention is characterized by use of anoxidizer tank having a cylindrical midsection which can be bonded to aspacecraft inner surface by a layer of elastomeric adhesive. Anelongated solid-fuel motor case is mechanically rigidly secured to acentral rear surface of the tank, and the case terminates in a throatand nozzle. The elastomeric-adhesive bonding of tank to spacecraft formsthe sole support for the rocket system, and separate support for themotor case is not required.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a side view of a hybrid rocket as mounted in a wingedspacecraft (shown in cross section);

FIG. 2 is an exploded view of components of the rocket;

FIG. 3 is a sectional perspective view of a forward end of a solid-fuelmotor case of the rocket;

FIG. 4 is a partial sectional view shown by the junction of the motorcase with an oxidizer tank; and

FIG. 5 is a partial sectional view of the oxidizer tank.

DESCRIPTION OF THE PREFERRED EMBODIMENT

FIG. 1 shows a hybrid rocket 10 as mounted in a winged suborbitalspacecraft fuselage 11, partially shown in cross section, and having apilot compartment 12. The major components of rocket 10 are an oxidizertank 14 with an elongated cylindrical midsection 15 secured to an innersurface 17 of the fuselage, the tank having a rear bulkhead (describedbelow) to which is secured a cylindrical motor 19 terminating in anozzle 20. A retractable landing gear and the spacecraft wing andempennage are omitted in FIG. 1 for clarity.

Oxidizer tank 14, shown in cross section in FIG. 5, has a relativelythin (e.g., 0.1-0.125 inch) inner liner 22 of an epoxy-fiberglassprepreg composite material. The inner liner is overwound with a layer 23of graphite-fiber tow coated with an epoxy matrix. The thickness of thetow/epoxy layer is about one-fourth inch in cylindrical midsection 15 ofthe tank, and gradually thickens toward the opposed tank ends to about1½ inch where the tank captures and is in sealed engagement withcylindrical forward and rear flanges 26 and 27 (FIG. 5).

Sealingly clamped to forward flange 26 is a forward bulkhead 29 whichcloses the front end of the tank, but provides a valved inlet (notshown) for loading of oxidizer. Similarly sealingly clamped to rearflange 27 is a forwardly convex rear bulkhead 30 which closes the rearend of the tank. An oxidant delivery valve and igniter system (notshown) are mounted on the rear bulkhead within the tank.

Tank cylindrical midsection 15 is covered by a cylindrical skirt 32which, in a preferred form, is made of about eleven plies or layers offlexible fiberglass cloth, the adhesively bound plies being oriented at45 degrees to the longitudinal axis of the tank. The inside diameter ofthe skirt is slightly larger than the outside diameter of the tankmidsection, enabling injection of a thick (about 0.1 inch) layer of anelastomeric adhesive PROSEAL P5890, Class B, made by PRC Desoto issatisfactory) to bond the skirt to the tank.

The thus-bonded skirt and tank make a slip fit within a mating innercylindrical surface 34 of a fuselage of spacecraft 11, enablinginjection of a thin (about 0.02 inch) layer of a rigid bonding adhesiveto secure the tank and skirt to the fuselage. An adhesive such as HYSOL9396 is satisfactory. The tank-skirt and skirt-fuselage bondings formthe sole support for rocket 10 within the spacecraft.

FIG. 4 is a partial cross-sectional view of the rigid junction of motor19 to oxidizer-tank rear flange 27 and rear bulkhead 30. Motor case 36has a silicon-phenolic inner liner 37 of about 0.1-inch thickness, andan outer wrap 38 of carbon-fiber tow of about 0.3-inch thickness. Solidfuel 39 is cast within the case, leaving open portions around the casecenterline. A compression-molded head insulation 40 is formed ahead ofthe solid fuel, and has a forwardly convex front end 41 following amating curvature of the rear end of bulkhead 30. The forward end of case36 is outwardly flared to seat against a cylindrical steel wedge 43 oftriangular cross section, and having a front surface 44 with an O-ringseal 45 seated against an outwardly extending cylindrical flange 46 onrear bulkhead 30.

An aluminum clamping ring 50 having a forwardly and outwardly taperedinner surface 51 fitted over the forward end of the motor case, andagainst the rear surface of flange 46. A number of threaded fasteners 52extend through mating openings in ring 50, flange 46, and flange 27 tonuts 53 to tightly clamp those components together, thereby rigidlysecuring the motor case to the tank. Additional O-ring seals 54 areseated in flange 46 and a forwardly extending ring 55 formed integrallywith the rear bulkhead. A perforated cylindrical tube 56 is secured toring 55, and extends forwardly within the oxidizer tank to provide aslosh baffle.

The oxidizer tank is loaded with nitrous oxide pressurized to about 700psi, at which pressure the oxidizer is a liquid. Outward swelling of thepressurized tank is at least partially absorbed by the thick layer ofelastomeric adhesive bonding the tank skirt to the fuselage, and thisadhesive layer also absorbs and damps motor vibrations.

Apart from cylindrical midsection 15, the oxidizer tank roughlyapproximates a sphere providing a pressure vessel strong enough tosupport the loads imposed by a fired motor. This strength permits themotor to be cantilevered off the rear end of the tank, and withoutfurther support from the fuselage. This single connection enables easyreplacement of the motor (typically a single-use component), as well asaccomodating motors of varying length, and reducing the number ofpossible leakage paths. Further, no additional weight is incurred inmotor support, as the strong tank/fuselage mounting supports the entirerocket system.

1. In a suborbital spacecraft having a fuselage with a generally central and open interior cylindrical surface, a hybrid-rocket propulsion system, comprising an oxidizer tank having an outer and central cylindrical surface covered by a skirt, and insertable within the fuselage interior surface, the skirt and tank being secured together by an elastomeric means, and an outer surface of the skirt being adhesively secured to the fuselage interior surface; an elongated and generally cylindrical solid-fuel motor case rigidly secured to a central rear surface of the oxidizer tank, the case having a rear end defining a throat, and a nozzle extending behind the fuselage; the case being cantilevered behind the tank, said oxidizer tank providing sufficient support to said motor case for entirely supporting said motor case.
 2. The propulsion system of claim 1, wherein the tank has an inner liner of an epoxy-fiberglass composite material, the inner liner being overwound with an outer layer of graphite-fiber tow and an epoxy matrix.
 3. The propulsion system of claim 2, wherein the skirt is fiberglass.
 4. The propulsion system of claim 3, wherein the elastomeric means is a layer of an elastomeric adhesive bonding the skirt to the tank.
 5. The propulsion system of claim 4 wherein the elastomeric adhesive has a thickness of about 0.1 inch.
 6. The propulsion system of claim 2, wherein the tank inner liner and outer layer are wound around and secured to forward and rear flanges, the flanges in turn being secured to forward and rear bulkheads respective, the bulkheads sealing the tank.
 7. The propulsion system of claim 6, wherein the motor case is rigidly secured to the rear bulkhead by releasable fasteners, enabling replacement of a fired motor case.
 8. The propulsion system of claim 1 wherein the elastomeric means and an adhesive adhesively securing the skirt to the fuselage form the sole support of the propulsion system to the fuselage.
 9. The propulsion system of claim 1 wherein said motor case is cantilevered from said oxidizer tank and is free from other attachment to the fuselage.
 10. The propulsion system of claim 1 wherein said motor case is only secured to said oxidizer tank.
 11. In a suborbital spacecraft having a fuselage with a generally central and open interior cylindrical surface, a hybrid-rocket propulsion system, comprising an oxidizer tank having an outer surface covered by a skirt, and insertable within the fuselage interior surface, the skirt and tank being secured together by an elastomeric means, and an outer surface of the skirt being adhesively secured to the fuselage interior surface; a solid-fuel motor case secured to a rear surface of the oxidizer tank, the case having a rear end defining a throat, and a nozzle; the case being cantilevered behind the tank, wherein said oxidizer tank provides sufficient support to said motor case for entirely supporting said motor case.
 12. The propulsion system of claim 11, wherein the elastomeric means is a layer of an elastomeric adhesive bonding the skirt to the tank.
 13. The propulsion system of claim 12 wherein the elastomeric adhesive has a thickness of about 0.1 inch.
 14. The propulsion system of claim 11, wherein the tank has an inner liner of an epoxy-fiberglass composite material, the inner liner being overwound with an outer layer of graphite-fiber tow and an epoxy matrix.
 15. The propulsion system of claim 14, wherein the skirt is fiberglass.
 16. The propulsion system of claim 14, wherein the tank inner liner and outer layer are wound around and secured to forward and rear flanges, the flanges in turn being secured to forward and rear bulkheads respective, the bulkheads sealing the tank.
 17. The propulsion system of claim 11, wherein the motor case is rigidly secured to the rear bulkhead by releasable fasteners, enabling replacement of a fired motor case.
 18. The propulsion system of claim 11 wherein the elastomeric means and an adhesive adhesively securing the skirt to the fuselage form the sole support of the propulsion system to the fuselage.
 19. The propulsion system of claim 11 wherein said motor case is entirely supported through said oxidizer tank. 